Aircraft identification and docking guidance systems

ABSTRACT

A laser range finder (LRF) is used to identify an aircraft approaching a gate. The LRF is directed at the aircraft, and from the echoes, a profile is derived and compared to known profiles. To distinguish among aircraft with similar profiles, the LRF is directed at a volume in which a feature such as an engine is expected and at another volume in which the engine is not expected. The echoes from the two volumes are used to determine whether the engine is in its expected location. If so, the aircraft is identified as the correct type and is allowed to dock at the gate. Otherwise, the aircraft is stopped. The nose height can be used as yet another identifying criterion.

REFERENCE TO RELATED APPLICATIONS

[0001] This is a continuation-in-part of U.S. patent application Ser.No. 09/429,609, filed Oct. 29, 1999, currently pending, which is acontinuation-in-part of U.S. patent application Ser. No. 08/817,368,filed Jul. 17, 1997, now U.S. Pat. No. 6,023,665, which is the U.S.national stage of PCT International Application No. PCT/SE94/00968,filed Oct. 14, 1994, published Apr. 25, 1996, as WO 96/12265 A1. Thedisclosures of the parent applications are hereby incorporated byreference in their entireties into the present disclosure.

BACKGROUND OF THE INVENTION

[0002] 1. Field of the Invention

[0003] This invention relates to systems for locating, identifying andtracking objects. More particularly, it related to aircraft location,identification and docking guidance systems and to ground trafficcontrol methods for locating and identifying objects on an airfield andfor safely and efficiently docking aircraft at such airport.

[0004] 2. Description of Related Art

[0005] In recent years there has been a significantly increased amountof passenger, cargo and other aircraft traffic including take offs,landings and other aircraft ground traffic. Also, there has been amarked increase in the number of ground support vehicles which arerequired to off load cargo, provide catering services and on goingmaintenance and support of all aircraft. With that substantial increasein ground traffic has come a need for greater control and safety in thedocking and identification of aircraft on an airfield.

[0006] Exemplary of prior art systems which have been proposed fordetecting the presence of aircraft and other traffic on an airfield arethose systems disclosed in U.S. Pat. No. 4,995,102; European Patent No.188 757; and PCT Published Applications WO 93/13104 and WO 93/15416.

[0007] However, none of those systems have been found to be satisfactoryfor detection of the presence of aircraft on an airfield, particularly,under adverse climatic conditions causing diminished visibility such asencountered under fog, snow or sleet conditions. Furthermore, none ofthe systems disclosed in the prior references are capable of identifyingand verifying the specific type of an approaching aircraft. Stillfurther, none of the prior systems provide adequate techniques fortracking and docking an aircraft at a designated stopping point such asan airport loading gate. Also, none of the prior systems have providedtechniques which enable adequate calibration of the instrument therein.

[0008] The system disclosed in the above-cited parent application seeksto overcome the above-noted problems though profile matching. Lightpulses from a laser range finder (LRF) are projected in angularcoordinates onto the airplane. The light pulses are reflected off theairplane to detect a shape of the airplane or of a portion of theairplane, e.g., the nose. The detected shape is compared with a profilecorresponding to the shape of a known model of airplane to determinewhether the detected shape corresponds to the shape of the known model.

[0009] However, that system has a drawback. Often, two or more models ofairplanes have nose profiles so similar that one model is oftenmisidentified as another. In particular, in adverse weather, many echoesare lost, so that profile discrimination becomes decreasingly reliable.Since the models are similar but not identical in body configuration, acorrect docking position for one can cause an engine on another to crashinto a physical obstacle.

[0010] Thus, it has been a continuing problem to provide systems whichare sufficiently safe and reliable over a wide range of atmosphericconditions to enable detection of objects such as aircraft and otherground traffic on an airfield.

[0011] In addition, there has been a long standing need for systemswhich are not only capable of detecting objects such as aircraft, butwhich also provide for the effective identification of the detectedobject and verification of the identity of such object, for example, adetected aircraft with the necessary degree of certainty regardless ofprevailing weather conditions and magnitude of ground traffic.

[0012] There has also been a long standing, unfulfilled need for systemswhich are capable of accurately and efficiently tracking and guidingobjects such as incoming aircraft to a suitable stopping point such asan airport loading gate. In addition, the provision of accurate andeffective calibration techniques for such systems has been a continuingproblem requiring resolution.

SUMMARY OF THE INVENTION

[0013] It will be readily apparent from the above that a need exists inthe art for a more accurate identification of aircraft.

[0014] It is therefore a primary object of the invention to distinguishamong multiple models of aircraft with identical or almost identicalnose shapes.

[0015] It is a further object of the invention to improve the detectionof aircraft so as to avoid accidents during aircraft docking.

[0016] To achieve the above and other objects, the present inventionidentifies aircraft in a two-step process. First, the profile matchingis performed as known from the above-identified parent application.Second, at least one aircraft criterion matching is performed. In theaircraft criterion matching, a component of the aircraft, such as theengine, is selected as a basis for distinguishing among aircraft. Thedisplacement of that component from another, easily located component,such as the nose, is determined in the following manner. An inner volumein which the engine is expected is defined, and an outer volumesurrounding the inner volume is also defined. The LRF is directed at theinner and outer volumes to produce echoes from both volumes. A ratio istaken of the number of echoes in the inner volumes to the number ofechoes in both volumes. If that echo exceeds a given threshold, theengine is determined to be present in the inner volume, and the aircraftis considered to be identified. If the identification of the aircraft isstill ambiguous, another aircraft criterion, such as the tail, can bedetected.

[0017] The aircraft criteria chosen for the second phase of theidentification are physical differences that can be detected by a laserrange finder. An example of such a criterion is the position, sidewaysand lengthwise, of an engine in relation to the aircraft nose. Toconsider an aircraft identified, the echo pattern must not only reflecta fuselage of correct shape. It must also reflect that there is anengine at a position, relative to the nose, where the expected aircraftdoes have an engine. Other examples of criteria that can be used are theposition of the main gear, the position of the wings and the position ofthe tail.

[0018] The matching is preferably done only against the criteriaspecific for the expected aircraft type. It would be very time consumingto match against the criteria of all other possible types. Such matchingwould have to be against every type of aircraft that may land at aspecific airport.

[0019] For each gate there is a defined a stopping position for eachaircraft type that is planned to dock at that gate. There might be asafety risk for any other type to approach the gate. The stoppingposition is defined so that there is a sufficient safety margin betweenthe gate and the aircraft to avoid collision. The stopping position foreach aircraft type is often defined as the position of the nose gearwhen the door is in appropriate position in relation to the gate. Thereis a database in the system where the distance from the nose to the nosegear is stored for each aircraft type. The docking system guides theaircraft with respect to its nose position and stops the aircraft withits nose in a position where the correct type will have its nose gear inthe correct stop position. If the wrong type is docked and if it has itswings or engines closer to the nose than the correct type, there is arisk of collision with the gate.

[0020] During the aircraft criteria phase, all aircraft criteriaspecified for the expected aircraft type can be checked. If an aircrafthas a profile that can be used to discriminate it from any other type,which is rarely the case, the profile will be the only aircraftcriterion. Otherwise, another criterion such as the position of theengine is checked, and if the identification is still ambiguous, stillanother criterion such as the position of the tail is checked.

[0021] The LRF is directed to obtain echoes from the inner and outervolumes. If the ratio of the number of echoes from within the innervolume to the number of echoes from within both volumes is larger than athreshold value, the aircraft is identified as having an engine at theright position, and that specific criterion is thus fulfilled. The ratioof the echo numbers is, however, just an example of a test used toevaluate the presence of an engine at the right position or to determinewhether the echoes come from some other source, e.g., a wing. In casesin which that is the only criterion, the aircraft is considered to beidentified. Otherwise, the other specified criteria (e.g., the height ofthe nose of the aircraft or evaluation of another aircraft criterion)have to be fulfilled.

[0022] If necessary, several characteristics, such as the tail, gears,etc., can be used to identify one specific type. The inner and outervolumes are then defined for each geometrical characteristic to be usedfor the identification. The exact extension of the volumes is dependenton the specific aircraft type and so is the threshold value.

[0023] A further identification criterion is the nose height. The noseheight is measured so as to allow the horizontal scan to be placed overthe tip of the nose. The measured nose height is also compared with theheight of the expected aircraft. If the two differ by more than 0.5 m,the aircraft is considered to be of wrong type, and the docking isstopped. The value 0.5 m is given by the fact that the ground heightoften varies along the path of the aircraft which makes it difficult tomeasure with higher accuracy.

[0024] The invention lends itself to the use of “smart” algorithms whichminimize the demand on the signal processing at the same time as theyminimize the effect of adverse weather and bad reflectivity of aircraftsurface. The advantage is that low-cost microcomputers can be used,and/or computer capacity is freed for other tasks, and that docking ispossible under almost all weather conditions.

[0025] One important algorithm in that respect is the algorithm forhandling of the reference profiles. The profile information is stored asa set of profiles. Each profile in the set reflects the expected echopattern for the aircraft at a certain distance from the system. Theposition of an aircraft is calculated by calculating the distancebetween the achieved echo pattern with the closest reference profile.The distance interval between the profiles in the set is chosen so shortthat the latter calculation can be made using approximations and stillmaintain the necessary accuracy. Instead of using scaling with a numberof multiplications, which is a demanding operation, simple addition andsubtraction can be used.

[0026] Another important algorithm is the algorithm for determining anaircraft's lateral deviation from its appropriate path. That algorithmuses mainly additions and subtractions and only very few multiplicationsand divisions. The calculation is based on areas between the referenceprofile and the echo pattern. As those areas are not so much affected byposition variations or absence of individual echoes the algorithmbecomes very insensitive to disturbances due to adverse weather.

[0027] The calibration procedure enables a calibration check against anobject at the side of the system. The advantage is that such acalibration check is possible also when no fixed object is available infront of the system. In most cases, there are no objects in front of thesystem that can be used. It is very important to make a calibrationcheck regularly. Something might happen to the system, e.g., such thatthe aiming direction of the system is changed. That can be due to anoptical or mechanical error inside the system or it can be due to amisalignment caused by an external force such as from a passing truck.If that happens, the system may guide an aircraft to a collision withobjects at the side of its appropriate path.

[0028] Another useful aspect of the present invention is that it caneasily be adapted to take into account the yaw angle of the aircraft.The yaw angle is useful to know for two reasons. First, knowledge of theyaw angle facilitates accurate docking of the aircraft. Second, once theyaw angle is determined, the profile is rotated accordingly, for moreaccurate matching.

[0029] In the verification process it is determined whether certaingeometric characteristics, such as an engine, are present in a certainposition, e.g., relative to the nose. If the aircraft is directed at anangle towards the docking guidance system (DGS), which is often thecase, that angle has to be known, in order to know where to look for thecharacteristics. The procedure is as follows:

[0030] 1. Convert the polar coordinates (angle, distance) of the echoesto Cartesian coordinates (x,y).

[0031] 2. Calculate the yaw angle.

[0032] 3. Rotate the echo profile to match the yaw angle calculated forthe aircraft.

[0033] 4. Determine the existence of the ID characteristics.

[0034] The yaw angle is typically calculated through a technique whichinvolves finding regression angles on both sides of the nose of theaircraft. More broadly, the geometry of the part of the aircraft justbehind the nose is used. Doing so was previously considered to beimpossible.

[0035] Still another aspect of the invention concerns the center linespainted in the docking ara. Curved docking center lines are painted asthe correct path for the nose wheel to follow, which is not the path forthe nose. If a DGS does not directly measure the actual position of thenose wheel, the yaw angle is needed to calculate it based on measureddata, such as the position of the nose. The position of the nose wheelin relation to the curved center line can then be calculated.

BRIEF DESCRIPTION OF THE DRAWINGS

[0036] The features and advantages of the invention will become apparentfrom the following detailed description taken in connection with theaccompanying drawings wherein:

[0037]FIG. 1 is a view illustrating the system as in use at an airport;

[0038]FIG. 2 is a diagrammatic view illustrating the general componentryof a preferred system in accordance with the present invention;

[0039]FIG. 3 is a top plan view illustrating the detection area in frontof a docking gate which is established for purposes of detection andidentification of approaching aircraft;

[0040]FIGS. 4A and 4B together show a flow chart illustrating the mainroutine and the docking mode of the system;

[0041]FIG. 5 is a flow chart illustrating the calibration mode of thesystem;

[0042]FIG. 6 is a view illustrating the components of the calibrationmode;

[0043]FIG. 7 is a flow chart illustrating the capture mode of thesystem;

[0044]FIG. 8 is a flow chart illustrating the tracking phase of thesystem;

[0045]FIG. 9 a is flow chart illustrating the height measuring the phaseof the system;

[0046]FIG. 10 is a flow chart illustrating the identification phase ofthe system.

[0047]FIG. 11 is a flow chart illustrating the aircraft criterion phaseof the system;

[0048]FIG. 12 is a diagram showing inner and outer volumes around anaircraft engine used in the aircraft criterion phase;

[0049]FIG. 13 is a diagram showing the tolerance limits of the measurednose-to-engine distance for accepting an aircraft into a gate;

[0050]FIG. 14 is a diagram showing the dependence of the safety marginon the nose-to-engine distance in a situation in which an aircraft ofthe wrong type is docked at the gate

[0051]FIG. 15 is a flow chart showing the basic steps used inrecognizing an aircraft which is at a yaw angle to the gate;

[0052]FIG. 15A is a diagram showing the geometry of the yaw angle;

[0053]FIG. 16 is a diagram showing the geometry used in determining theregression lines which are used in calculating the yaw angle;

[0054]FIG. 17 is a flow chart showing the steps used in calculating theyaw angle;

[0055]FIG. 18 is a diagram showing the geometry used in rotating an echoprofile;

[0056]FIG. 19 is a flow chart showing the steps used in rotating theecho profile;

[0057]FIG. 20 is a flow chart showing the steps used in calculating anoffset of a nose wheel of an aircraft from a center line;

[0058]FIG. 21 is a diagram showing the geometry of the position of thenose wheel relative to that of the nose; and

[0059]FIG. 22 is a diagram showing the geometry of the position of thenose wheel relative to the center line.

[0060] Table I is a preferred embodiment of a Horizontal ReferenceProfile Table which is employed to establish the identity of an aircraftin the systems of the present invention;

[0061] Table II is a preferred embodiment of a Comparison Table which isemployed in the systems of the present invention for purposes ofeffectively and efficiently docking an aircraft.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

[0062] Reference is now made to FIGS. 1-22 and Tables I-II, in whichlike numerals designate like elements throughout the several views.Throughout the following detailed description, numbered stages depictedin the illustrated flow diagrams are generally indicated by elementnumber in parenthesis following such references.

[0063] Referring to FIG. 1, the docking guidance systems of the presentinvention generally designated 10 in the drawings provide for thecomputerized location of an object, verification of the identity of theobject and tracking of the object, the object preferably being anaircraft 12. In operation, once the control tower 14 lands an aircraft12, it informs the system that a plan is approaching gate 16 and thetype of aircraft (i.e., 747, L-1011, etc.) expected. The system 10 thenscans the area in front of the gate 16 until it locates an object thatit identifies as an airplane 12. The system 10 then compares themeasured profile of the aircraft 12 with a reference profile for theexpected type of aircraft and evaluates other geometric criteriacharacteristic of the expected aircraft type. If the located aircraftdoes not match the expected profile and the other criteria, the systeminforms or signals the tower 14 and shuts down.

[0064] If the object is the expected aircraft 12, the system 10 tracksit into the gate 16 by displaying in real time to the pilot the distanceremaining to the proper stopping point 29 and the lateral position 31 ofthe plane 12. The lateral position 31 of the plane 12 is provided on adisplay 18 allowing the pilot to correct the position of the plane toapproach the gate 16 from the correct angle. Once the airplane 12 is atits stopping point 53, that fact is shown on the display 18 and thepilot stops the plane.

[0065] Referring to FIG. 2, the system 10 includes a Laser Range Finder(LRF) 20, two mirrors 21, 22, a display unit 18, two step motors 24, 25,and a microprocessor 26. Suitable LRF products for use herein are soldby Laser Atlanta Corporation and are capable of emitting laser pulsesand receiving the reflections of those pulses reflected off of distantobjects and computing the distance to those objects.

[0066] The system 10 is arranged such that there is a connection 28between the serial port of the LRF 20 and the microprocessor 26. Throughthat connection, the LRF 20 sends measurement data approximately every{fraction (1/400)}th of a second to the microprocessor 26. The hardwarecomponents generally designated 23 of the system 20 are controlled bythe programmed microprocessor 26. In addition, the microprocessor 26feeds data to the display 18. As the interface to the pilot, the displayunit 18 is placed above the gate 16 to show the pilot how far the planeis from its stopping point 29, the type of aircraft 30 the systembelieves is approaching and the lateral location of the plane 31. Usingthat display, the pilot can adjust the approach of the plane 12 to thegate 16 to ensure the plane is on the correct angle to reach the gate.If the display 18 shows the wrong aircraft type 30, the pilot can abortthe approach before any damage is done. That double check ensures thesafety of the passengers, plane and airport facilities because if thesystem tries to dock a larger 747 at a gate where a 737 is expected, itlikely will cause extensive damage.

[0067] In addition to the display 18, the microprocessor 26 processesthe data from LRF 20 and controls the direction of the laser 20 throughits connection 32 to the step motors 24, 25. The step motors 24, 25 areconnected to the mirrors 21, 22 and move them in response toinstructions from the microprocessor 26. Thus, by controlling the stepmotors 24, 25, the microprocessor 26 can change the angle of the mirrors21, 22 and aim the laser pulses from the LRF 20.

[0068] The mirrors 21, 22 aim the laser by reflecting the laser pulsesoutward over the tarmac of the airport. In the preferred embodiment, theLRF 20 does not move. The scanning by the laser is done with mirrors.One mirror 22 controls the horizontal angle of the laser while the othermirror 21 controls the vertical angle. By activating the step motors 24,25, the microprocessor 26 controls the angle of the mirrors and thus thedirection of the laser pulse.

[0069] The system 10 controls the horizontal mirror 22 to achieve acontinuous horizontal scanning within a ±10 degree angle inapproximately 0.1 degree angular steps which are equivalent to 16microsteps per step with the Escap EDM-453 step motor. One angular stepis taken for each reply from the reading unit, i.e., approximately every2.5 ms. The vertical mirror 21 can be controlled to achieve a verticalscan between ±20 and −30 degrees in approximately 0.1 degree angularsteps with one step every 2.5 ms. The vertical mirror is used to scanvertically when the nose height is being determined and when theaircraft 12 is being identified. During the tracking mode, the verticalmirror 21 is continuously adjusted to keep the horizontal scan trackingthe nose tip of the aircraft 12.

[0070] Referring to FIG. 3, the system 10 divides the field in front ofit by distance into three parts. The farthest section, from about 50meters out, is the capture zone 50. In that zone 50, the system 10detects the aircraft's nose and makes a rough estimate of lateral andlongitudinal position of the aircraft 12. Inside the capture zone 50 isthe identification area 51. In that area, the system 10 checks theprofile of the aircraft 12 against a stored profile 51. In that area,the system 10 checks the profile of the aircraft 12 in that region,related to a predetermined line, on the display 18. Finally, nearest tothe LRF 20 is the display or tracking area 52. In the display area 52,the system 10 displays the lateral and longitudinal position of theaircraft 12 relative to the correct stopping position with its highestdegree of accuracy. At the end of the display area 52 is the stoppingpoint 53. At the stopping point 53, the aircraft will be in the correctposition at the gate 16.

[0071] In addition to the hardware and software, the system 10 maintainsa database containing reference profiles for any type of aircraft itmight encounter. Within that database, the system stores the profile foreach aircraft type as a horizontal and vertical profile reflecting theexpected echo pattern for that type of aircraft.

[0072] Referring to Table I, the system maintains the horizontal profilein the form of a Table I whose rows 40 are indexed by angular step andwhose columns 41 are indexed by distance from the stopping position forthat type of aircraft. In addition to the indexed rows, the tablecontains a row 42 providing the vertical angle to the nose of the planeat each distance from the LRF a row 44 providing the form factor, k, forthe profile and a row 45 providing the number of profile values for eachprofile distance. The body 43 of the Table I contains expected distancesfor that type of aircraft at various scanning angles and distances fromthe stopping point 53.

[0073] Theoretically, the 50 angular steps and the 50 distances to thestopping point 53 would require a Table I containing 50×50, or 2500,entries. However, Table I will actually contain far fewer entriesbecause the profile will not expect a return from all angles at alldistances. It is expected that a typical table will actually containbetween 500 and 1000 values. Well known programming techniques providemethods of maintaining a partially full table without using the memoryrequired by a full table.

[0074] In addition to the horizontal profile, the system 10 maintains avertical profile of each type of aircraft. That profile is stored in thesame manner as the horizontal profile, except that its rows are indexedby angular steps in the vertical direction and its column index containsfewer distances from the stopping position than the horizontal profile.The vertical profile requires fewer columns because it is used only foridentifying the aircraft 12 and for determining its nose height, whichtake place at a defined range of distances from the LRF 20 in theidentification area 51. Consequently, the vertical profile stores onlythe expected echoes in that range without wasting data storage space onunneeded values.

[0075] The system 10 uses the previously described hardware and databaseto locate, identify and track aircraft using the following procedures:

[0076] Referring to FIGS. 4A and 4B, the software running on themicroprocessor performs a main routine containing subroutines for thecalibration mode 60, capture mode 62 and docking mode 400. Themicroprocessor first performs the calibration mode 60, then the capturemode 62 and then the docking mode 400. Once the aircraft 12 is docked,the program finishes. Those modes are described in greater detail asfollows:

[0077] Calibration Mode

[0078] To ensure system accuracy, the microprocessor 26 is programmed tocalibrate itself in accordance with the procedure illustrated in FIG. 5before capturing an aircraft 12 and at various intervals duringtracking. Calibrating the system 10 ensures that the relationshipbetween the step motors 24, 25 and the aiming direction is known. Thelength measuring ability of the LRF 20 is also checked.

[0079] Referring to FIG. 6, for calibration, the system 10 uses a squareplate 66 with a known position. The plate 66 is mounted 6 meters fromthe LRF 20 and at the same height as the LRF 20.

[0080] To calibrate, the system sets (α,β) to (0,0), causing the laserto be directed straight forward. The vertical mirror 22 is then tiltedsuch that the laser beam is directed backwards to a rear or extra mirror68 which redirects the beam to the calibration plate 66. (100) Themicroprocessor 26 then uses the step motors 24, 25, to move the mirrors21, 22 until it finds the center of the calibration plate 66. Once itfinds the center of the calibration plate 66, the microprocessor 26stores the angles (α_(cp),β_(cp)) at that point and compares them tostored expected angles. (102) The system 10 also compares the reporteddistance to the plate 66 center with a stored expected value. (102) Ifthe reported values do not match the stored values, the microprocessor26 changes the calibration constants, which determine the expectedvalues, until they do. (104, 106) However, if any of those valuesdeviate too much from the values stored at installation, an alarm isgiven. (108)

[0081] Capture Mode

[0082] Initially, the airport tower 14 notifies the system 10 to expectan incoming airplane 12 and the type of airplane to expect. That signalputs the software into a capture mode 62 as outlined in FIG. 7. Incapture mode 62, the microprocessor 26 uses the step motors 24, 25 todirect the laser to scan the capture zone 50 horizontally for the plane12. That horizontal scan is done at a vertical angle corresponding tothe height of the nose of the expected type of aircraft at the midpointof the capture zone 50.

[0083] To determine the correct height to scan, the microprocessor 26computes the vertical angle for the laser pulse as:

α_(f) =arctan[(H−h)/l_(f)]

[0084] where H=the height of the LRF 20 above the ground, h=the noseheight of the expected aircraft, and l_(f)=the distance from the LRF 20to the middle of the capture zone 50. That equation results in avertical angle for the mirror 21 that will enable the search to be atthe correct height at the middle of the capture zone 50 for the expectedairplane 12.

[0085] Alternatively, the system 10 can store in the database values forβ_(f) for different types of aircraft at a certain distance. However,storing β_(f) limits the flexibility of the system 10 because it cancapture an aircraft 12 only a single distance from the LRF 20

[0086] In the capture zone 50 and using that vertical angle, themicroprocessor 26 directs the laser to scan horizontally in pulsesapproximately 0.1 degree apart. The microprocessor 26 scans horizontallyby varying α, the horizontal angle from a center line starting from theLRF-20, between ±α_(max), a value defined at installation. Typically,α_(max) is set to 50 which, using 0.1 degree pulses, is equivalent to 5degrees and results in a 10 degree scan.

[0087] The release of the laser pulses results in echoes or reflectionsfrom objects in the capture zone 50. The detection device of the LRF 20captures the reflected pulses, computes the distance to the object fromthe time between pulse transmission and receipt of the echo, and sendsthe calculated distance value for each echo to the microprocessor 26.The micro processor 26 stores, in separate registers in a data storagedevice, the total number of echoes or hits in each I degree sector ofthe capture zone 50. (70) Because the pulses are generated in 0.1 degreeintervals, up to ten echoes can occur in each sector. The microprocessor26 stores those hits in variables entitled s_(α) where a varies from 1to 10 to reflect each one degree slice of the ten degree capture zone50.

[0088] In addition to storing the number of hits per sector, themicroprocessor 26 stores, again in a data storage device, the distancefrom the LRF 20 to the object for each hit or echo. Storing the distanceto each reflection requires a storage medium large enough to store up toten hits in each 1 degree of the capture zone 50 or up to 100 possiblevalues. Because, in many cases, most of the entries will be empty, wellknown programming techniques an reduce those storage requirements belowhaving 100 registers always allocated for those values.

[0089] Once that data is available for a scan, the microprocessor 26computes the total number of echoes, S_(T), in the scan by summing thes_(α)'s. The microprocessor 26 then computes S_(M), the largest sum ofechoes in three adjacent sectors. (72) In other words, S_(M) is thelargest sum of (S_(α−1), S_(α), S_(α+1)).

[0090] Once it computes S_(M) and S_(T), the microprocessor 26determines whether the echoes are from an incoming airplane 12. If S_(M)is not greater than 24, no airplane 12 has been found and themicroprocessor 26 returns to the beginning of the capture mode 62. Ifthe largest sum of echoes, S_(M) is greater than 24 (74), a “possible”airplane 12 has been located. If a “possible” airplane 12 has beenlocated, the microprocessor checks if S_(M)/S_(T) is greater than 0.5(76), or the three adjacent sectors with the largest sum contain atleast half of all the echoes received during the scan.

[0091] If S_(M)/S_(M) is greater than 0.5, the microprocessor 26calculates the location of the center of the echo. (78, 82) The angularlocation of the center of the echo is calculated as:

α₁=α_(v)+(S _(α+1) −S _(α−1))/(S _(α−2) +S _(α) +S _(α+1))

[0092] where S_(α) is the S_(α) that gave S_(M) and α_(v) is the angularsector that corresponds to that S_(α).

[0093] The longitudinal position of the center of the echo is calculatedas $l_{t} = {\frac{1}{n}{\sum\limits_{i = 1}^{10}\quad l_{avi}}}$

[0094] where the l_(avi) are the measured values, or distances to theobject, for the pulses that returned an echo from the sector α_(v) andwhere n is the total number of measured values in that sector. (78, 82)Because the largest possible number of measured values is ten, n must beless than or equal to ten.

[0095] However, if S _(M)/S_(T)<0.5, the echoes may have been caused bysnow or other aircraft at close range. If the cause is an aircraft atclose range, that aircraft is probably positioned fairly close to thecenterline so it is assumed that at should be zero instead of the abovecalculated value and that l_(t) should be the mean distance given by thethree middle sectors. (80) If the distance distribution is too large,the microprocessor 26 has not found an airplane 12 and it returns to thebeginning of the capture mode 62. (81).

[0096] After calculating the position of the aircraft 12, the system 10switches to docking mode 400.

[0097] Docking Mode

[0098] The docking mode 400 illustrated in FIGS. 4A and 4B includes fourphases, the tracking phase 84, the height measuring phase 86, theprofile recognition phase 404, and the aircraft criteria phase 408. Inthe tracking phase 84, the system 10 monitors the position of theincoming aircraft 12 and provides the pilot with information about axiallocation 31 and distance from the stopping point 53 of the plane throughthe display 18. The system 10 begins tracking the aircraft 12 byscanning horizontally.

[0099] Referring to FIG. 8, during the first scan in tracing phase, themicroprocessor 26 directs the LRF 20 to send out laser pulses in singleangular steps, αor, preferably, at 0.1 degree intervals between(α_(t)−α_(p)−10) and (α_(t)+α_(p)+10), where α_(t) is determined duringthe capture mode 62 as the angular position of the echo center and α_(p)is the largest angular position in the current profile column thatcontains distance values.

[0100] After the first scan, αis stepped back and forth with one stepper received LRF value between (α_(s)−α_(p)−10) and (α_(s) +α_(p)+10),where α_(s) is the angular position of the azimuth determined during theprevious scan.

[0101] During the tracking phase 84, the vertical angle β is set to thelevel required for the identified craft 12 at its current distance fromthe LRF 20 which is obtained from the reference profile Table I. Thecurrent profile column is the column representing a position less thanbut closer to l_(t).

[0102] The microprocessor 26 uses the distance from the stopping point53 to find the vertical angle for the airplane's current distance on theprofile Table I. During the first scan, the distance, l_(t), calculatedduring the capture mode 62, determines the appropriate column of theprofile Table I and thus the angle to the aircraft 12. For eachsubsequent scan, the microprocessor 26 uses the β in the column of theprofile Table I reflecting the present distance from the stopping point53. (112)

[0103] Using the data from the scans and the data on the horizontalprofile Table I, the microprocessor 26 creates a Comparison Table II.The Comparison Table II is a two dimensional table with the number ofthe pulse, or angular step number, as the index 91, i, to the rows.Using that index, the following information, represented as columns ofthe table, can be accessed for each row: l_(i) 92, the measured distanceto the object on that angular step; l_(k) 93, the measured valuecompensated for the skew caused by the displacement (equal to l_(i)minus the quality s_(m), the total displacement during the last scan,minus the quality i times s_(p), the average displacement during eachstep in the last scan, i.e., l_(i)=(s_(m)−is_(p))); d_(i) 94, thedistance between the generated profile and the reference profile (equalto r_(ij), the profile value for the corresponding angle at the profiledistance j minus I_(ki)); a_(l) 95, the distance the nose of theaircraft and the measuring equipment (equal to r_(j50), the referenceprofile value at zero degrees, minus d_(i)); a_(e) 96, the estimatednose distance after each step (equal to a_(m), the nose distance at theend of the last scan, minus the quantity i times s_(p)); a_(d), thedifference between the estimated and measured nose distance (equal tothe absolute value of a_(i) minus a_(c)); and Note 97 which indicatesthe echoes that are likely caused by an aircraft.

[0104] During the first scan in the tracking phase 84, the system 10uses the horizontal profile column representing an aircraft position, j,less than but closest to the value of l_(t). For each new scan, theprofile column whose value is less than but closest to (a_(m)−s_(m)) ischosen where a_(m) is the last measured distance to the aircraft 12 ands_(m) is the aircraft's displacement during the last scan. Additionally,the values of the profile are shifted sideways by α_(s) to compensatefor the lateral position of the aircraft. (112)

[0105] During each scan, the microprocessor 26 also generates a DistanceDistribution Table (DDT). That table contains the distribution of as_(i) value as they appear in the Comparison Table II. Thus, the DDT hasan entry representing the number of occurrences of each value of a_(i)in the Comparison Table II in 1 meter increments between 10 to 100meters.

[0106] After every scan, the system 10 uses the DDT to calculate theaverage distance a_(m), to the correct stopping point 53. Themicroprocessor 26 scans the data in the DDT to find the two adjacententries in the DDT for which the sum of their values is the largest. Themicroprocessor 26 then flags the Note 97 column in the Comparison TableII for each row containing an entry for a_(i) corresponding to either ofthe two DDT rows having the largest sum. (114)

[0107] The system 10 then determines the lateral deviation of offset.(116) The microprocessor 26 first sets:

2d=α _(max) −α _(min)

[0108] where α_(max) and α_(min) are the highest and lowest a values fora continuous flagged block of d_(j) values in the Comparison Table II.Additionally, the microprocessor 26 calculates:

Y₁=Σd_(i)

[0109] for the upper half of the flagged d_(j) in the block and:

Y2=Σd_(i)

[0110] for the lower half of the block. Using Y₁ and Y₂, “a” 116 iscalculated as:

a-kx(Y ₁ −Y ₂)/d²

[0111] where k is given in the reference profile. If “a” exceeds a givenvalue, preferably set to one, it is assumed that there is a lateraldeviation approximately equal to “a”. The l_(i) column of the ComparisonTable II is then shifted “a” steps and the Comparison Table II isrecalculated. The process continues until “a” is smaller than anempirically established value, preferably one. The total shift, α_(s),of the l_(i) column is considered equal to the lateral deviation oroffset. (116) If the lateral offset is larger than a predeterminedvalue, preferably set to one, the profile is adjusted sideways beforethe next scan. (118, 120)

[0112] After the lateral offset is checked, the microprocessor 26provides the total sideways adjustment of the profile, which correspondsto the lateral position 31 of the aircraft 12, on the display 18. (122)

[0113] The microprocessor 26 next calculates the distance to the nose ofthe aircraft, a_(m)

a _(m)=Σ(flagged a _(i))/N

[0114] where N is the total number of flagged a_(i). From a_(m), themicroprocessor 26 can calculate the distance from the plane 12 to thestopping point 53 by subtracting the distance from the LRF 20 to thestopping point 53 from the distance of the nose of the aircraft. (124)

[0115] Once it calculates the distance to the stopping point 53, themicroprocessor 26 calculates the average displacement during the lastscan, s_(m). The displacement during the last scan is calculated as:

S _(m) =a _(m−1) −a _(m)

[0116] where a_(m−1) and a_(m) belong to the last two scans. For thefirst scan in tracking phase 84, S_(m) is set to 0.

[0117] The average displacement during each step is calculated as:

S _(p) =S _(m) /P

[0118] where P is the total number of steps for the last scan cycle.

[0119] The microprocessor 26 will inform the pilot of the distance tothe stopping position 53 by displaying it on the display unit 18, 29. Bydisplaying the distance to the stopping position 29, 53 after each scan,the pilot receives constantly updated information in real time about howfar the plane 12 is from stopping.

[0120] If the aircraft 12 is in the display area 52, both the lateral 31and the longitudinal position 29 are provided on the display 18. (126,128) Once the microprocessor 26 displays the position of the aircraft12, the tracking phase ends.

[0121] Once it completes the tracking phase, the microprocessor 26verifies that tracking has not been lost by checking that the totalnumber of rows flagged divided by the total number of measured values,or echoes, in the last scan is greater than 0.5. (83) In other words, ifmore that 50% of the echoes do not correspond to the reference profile,tracking is lost. If tracking is lost and the aircraft 12 is greaterthan 12 meters from the stopping point, the system 10 returns to thecapture mode 62. (85) If tracking is lost and the aircraft 12 is lessthan or equal to 12 meters from the stopping point 53, the system 10turns on the stop sign to inform the pilot that it has lost tracking.(85, 87)

[0122] If tracking is not lost, the microprocessor 26 determines if thenose height has been determined. (13) If the height has not yet beendetermined, the microprocessor 26 enters the height measuring phase 86.If the height has already been determined, the microprocessor 26 checksto see if the profile has been determined (402).

[0123] In the height measuring phase, illustrated in FIG. 9, themicroprocessor 26 determines the nose height by directing the LRF 20 toscan vertically. The nose height is used by the system to ensure thatthe horizontal scans are made across the tip of the nose.

[0124] To check the nose height, the microprocessor 26 sets β to apredetermined value β_(max) and then steps it down in 0.1 degreeintervals once per received/reflected pulse until it reaches β_(min),another predetermined value. β_(min) and β_(max) are set duringinstallation and typically are −20 and 30 degrees respectively. After βreaches β_(min) the microprocessor 26 directs the step motors, 24, 25 upuntil it reaches β_(max). That vertical scanning is done with α set toα_(s), the azimuth position of the previous scan.

[0125] Using the measured aircraft distance, the microprocessor 26selects the column in the vertical profile table closest to the measureddistance. (140) Using the data from the scan and the data on thevertical profile table, the microprocessor 26 creates a comparison tableshown herein as Table II. Table II is a two dimensional table with thenumber of the pulse, or angular step number, as an index 91, i, to therows. Using that index, the following information, represented ascolumns of the table, can be accessed for each row: l_(i) 92, themeasured distance to the object on that angular step, l_(ki) 93, themeasured value compensated for the skew caused by the displacement(equal to l_(i) minus the quantity S_(m), the total displacement duringthe last scan, minus the quantity i times s_(p), the averagedisplacement during each step in the last scan), d_(i) 94, the distancebetween the generated profile and the reference profile (equal tor_(ij), the profile value for the corresponding angle at the profiledistance j, minus l_(ki)), a_(i) 95, the distance between the nose ofthe aircraft and the measuring equipment equal to r_(j50), the referenceprofile value at zero degrees, minus d_(i)), a_(e) 96, the estimatednose distance after each step (equal to a_(m), the nose distance at theend of the last scan, minus the quantity i times s_(p)), a_(d), thedifference between the estimated and measured nose distance (equal tothe absolute value of a_(i) minus a_(e)), and Note 97 which indicatesechoes that are likely caused by an aircraft 12.

[0126] During each scan the microprocessor 26 also generates a DistanceDistribution Table (DDT). That table contains the distribution of a_(i)values as they appear in Table II. Thus, the DDT has an entryrepresenting the number of occurrences of each value of a_(i) in TableII in I meter increments between 10 to 100 meters.

[0127] After every scan, the system 10 uses the DDT to calculate theaverage distance, a_(m), to the correct stopping point 53. Themicroprocessor 26 scans the data in the DDT to find the two adjacententries in the DDT for which the sum of their values is the largest. Themicroprocessor 26 then flags the Note 97 column in Table II for each rowcontaining an entry for a_(i) corresponding to either of the two DDTrows having the largest sum. (142)

[0128] Once it completes the calculation of the average distance to thecorrect stopping point 53, the microprocessor 26 calculates the averagedisplacement during the last scan, s_(m). The displacement during thelast scan is calculated as:

s _(m) a _(m−1) −a _(m)

[0129] where a_(m−1) and a_(m) belong to the last two scans. For thefirst scan in tracking phase 84, s_(m) is set to 0. The averagedisplacement s_(p) during each step is calculated as:

s _(p) =s _(m) /P

[0130] where P is the total number of steps for the last scan cycle.

[0131] Calculating the actual nose height is done by adding the nominalnose height, predetermined height of the expected aircraft when empty,to the vertical or height deviation. Consequently, to determine the noseheight, the system 10 first determines the vertical or height deviation.(144) Vertical deviation is calculated by setting:

2d=α _(max)−β_(min)

[0132] where β_(max) and β_(min) are the highest and lowest p value fora continuous flagged block of d_(i) values in the Comparison Table II.Additionally, the microprocessor 26 calculates:

Y₁=Σd_(i)

[0133] for the upper half of the flagged d_(i) in the block and;

Y₂=Σd_(i)

[0134] for the lower half of the block. Using Y₁ and Y₂, “a” iscalculated as

a=kx(Y ₁ −Y ₂)/d ²

[0135] where k is given in the reference profile. If “a” exceeds a givenvalue, preferably one, it is assumed that there is a vertical deviationapproximately equal to “a”. The 1; column is then shifted “a” steps, theComparison Table II is re-screened and “a” recalculated. That processcontinues until “a” is smaller than the given value, preferably one. Thetotal shift, β_(s), of the l_(i) column is considered equal to theheight deviation. (144) The β_(i) values in the vertical ComparisonTable II are then adjusted as β_(j)+Δβ_(j) where the height deviationΔβ_(j) is:

Δβ_(j)=β_(s)×(a _(mβ) +a _(s))/(a _(j) +a _(s))

[0136] and where a_(mβ) is the valid a_(m) value when β_(s) wascalculated.

[0137] Once the height deviation is determined, the microprocessor 26checks if it is bigger than a predetermined value, preferably one. (146)If the deviation is larger than that value, the microprocessor 26adjusts the profile vertically corresponding to that offset. (148) Themicroprocessor 26 stores the vertical adjustment as the deviation fromthe nominal nose height. (150) The actual height of the aircraft is thenominal nose height plus the deviation.

[0138] If the nose height is determined, or once the height measuringphase 86 is run, the microprocessor 26 enters the identification phaseillustrated in FIG. 10. (133, 88) In the identification phase 88, themicroprocessor 26 creates a Comparison Table II to reflect the resultsof another vertical scan and the contents of the profile table. (152,154). Another vertical scan is performed in the identification phase 88because the previous scan may have provided sufficient data for heightdetermination but not enough for identification. In fact, several scansmay need to be done before a positive identification can be made. Aftercalculating the vertical offset 156, checking that it is not too large(158) and adjusting the profile vertically corresponding to the offset(160) until the offset drops below a given amount, preferably one, themicroprocessor 26 calculates the average distance between marked echoesand the profile and the mean distance between the marked echoes and thataverage distance. (162)

[0139] The average distance d_(m) between the measured and correctedprofile and the deviation T from that average distance are calculatedafter vertical and horizontal scans as follows:

d _(m) =Σd _(i) /N

T−Σ|d _(i) −d _(m) |/N

[0140] If T is less than a given value, preferably 5, for both profiles,the aircraft 12 is judged to be of the correct type provided that asufficient number of echoes are received. (164) Whether a sufficientnumber of echoes is received is based on:

N/size >0.75

[0141] where N is the number of “accepted” echoes and “size” is themaximum number of values possible. If the aircraft 12 is not of thecorrect type, the microprocessor turns on the stop sign 136 and suspendsthe docking mode 400.

[0142] If the profile is determined (402), or once the profiledetermination phase is run (404), the microprocessor 26 determineswhether the aircraft criterion is determined (406). If not, the aircraftcriterion phase 408, which is illustrated in FIGS. 11 and 12, is run.

[0143] In order for the criterion to be fulfilled, echoes must bereturned from the location where there is an engine on the expectedaircraft. As there is some measurement uncertainty, there might beechoes that actually come from the engine but appear to come fromoutside the engine. Therefore, there must be defined a space Vi, calledthe inner volume or the active volume, around the engine, such thatechoes from within Vi are considered to come from the engine. FIG. 12shows a sample Vi around an engine 13 of an airplane 12.

[0144] An engine is characterized in that for a horizontal scan there isa reflecting surface surrounded by free space. In order to be able todiscriminate between an engine and, e.g., a wing, there must be definedanother space Vo around the engine where there must be no or very fewechoes. The space Vo is called the outer volume or the passive volume.FIG. 12 also shows a sample Vo around Vi.

[0145] The engine is defined by its coordinates (dx, dy, dz) for thecenter of the engine front relative to the nose and by its diameter D.Those parameters are stored in a database for all aircraft types.

[0146] Vi and Vo are defined by the extension sideways (x-direction) andlengthwise (z-direction) from that engine center. The vertical positionof the engine is given as (nose height+dy).

[0147] For an engine on the wing, Vi and Vo are defined by the followingranges of coordinates: Vi: x-direction: ±(D/2 + 1 m) z-direction: +3 m,−1 m Vo: x-direction: ±2 m from Vi z-direction: ±1.5 m from Vi

[0148] For tail engines the definition is the same except for Vo in thex-direction, which is given by +2 m from Vi. Otherwise echoes from thefuselage could fall within Vo and the criterion would not be fulfilled.

[0149] Finally, the criterion is

Vi/(Vi+Vo)>0.7

[0150] The threshold value 0.7 in the criterion is determinedempirically. So are the limits given above for Vi and Vo. At the momentthose values are chosen so that unnecessary ID failures are avoided andthey are different only dependent on if the engine is on the wing or onthe tail. As docking data is accumulated they will be adjusted, probablydifferent for different aircraft types, to achieve better and betterdiscrimination.

[0151] The aircraft criteria phase 408 applies the above principles asshown in the flow chart of FIG. 11. When the aircraft criteria phasestarts, the LRF is directed toward the engine or other selected aircraftcriterion in step 1102. In step 1104, the number of echoes in Vi isfound, and in step 1106, the number of echoes in Vo is found. In step1108, it is determined whether Vi/(Vi+Vo) exceeds the threshold value.If so, the aircraft criterion is indicated as met (OK) in step 1110.Otherwise, the aircraft criterion is indicated as unmet (not OK) in step1112.

[0152] If the aircraft criterion has been determined (406), or once theaircraft criterion phase is complete (408), the microprocessor 26determines whether the aircraft 12 has been identified. (410). If theaircraft 12 has been identified, the microprocessor 26 checks whetherthe aircraft 12 has reached the stop position. (412). If the stopposition is reached, the microprocessor 26 turns on the stop sign,whereupon the system 10 has completed the docking mode 400. (414) If theaircraft 12 has not reached the stop position, the microprocessor 26returns to the tracking phase 84.

[0153] If the aircraft 12 is not identified, the microprocessor 26checks whether the aircraft 12 is less than or equal to 12 meters fromthe stopping position 53. (416) If the aircraft 12 is not more than 12meters from the stopping position 53, the system 10 turns on the stopsign to inform the pilot that the identification has failed. (418) Afterdisplaying the stop sign, the system 10 shuts down.

[0154] If the aircraft 12 is more than 12 meters from the stopping point53, the microprocessor 26 returns to the tracking phase 84.

[0155] In one possible implementation, the nominal distance(longitudinal and lateral) from the nose to the engine is used as theaircraft criterion. In that implementation, docking is stopped if thenose-to-engine distance, as measured in step 408, is more than twometers shorter than that for the expected aircraft. If the difference iswithin two meters, it may still be possible to accept an aircraft of thewrong type safely. In the latter case, if the safety margin between theengine and a structure of the airport gate is three meters for thecorrect type of aircraft, the safety margin for the other type ofaircraft is still at least one meter. Tests have shown that the engineposition can be located to within about ±1 meter and that the noseheight can be determined to within ±0.5 meter.

[0156]FIG. 13 shows the nominal nose to engine distance of an aircraft12. The distance from the aircraft's nose to its engine 13 is ofparticular concern, since the engine 13 is in such a position thatmisidentification can result in a collision between the engine 13 and acomponent of the gate. Also shown are forward and backward tolerancelimits for the position of the engine 13 that define the forward andbackward extents of Vi.

[0157]FIG. 14 shows an application of the identification proceduredescribed above and in particular shows what may happen if the system isset up for a selected aircraft 12A, but another aircraft 12B attempts todock at that gate. If a type of aircraft 12B different from the selectedaircraft 12A is accepted into the gate, the aircraft 12B will be stoppedwith the nose in the same position in which the nose of the selectedaircraft 12A would be stopped. As a result, the safety margin, which isthe distance from the engine to the closest component of the gate, suchas the bridge 15, is different between the aircraft 12A and 12B if thenose-to-engine distances of those aircraft are different. As can be seenfrom FIG. 14, the safety margin for the aircraft 12B is equal to thesafety margin for the aircraft 12A minus the difference innose-to-engine distances. If, for example, the safety margin for theaircraft 12A is 3 m, and the nose-to-engine distance for the aircraft12B is 3.5 m shorter than that for the aircraft 12A, the engine 13B ofthe aircraft 12B will collide with the bridge 15. Therefore, if allaircraft types for which the nose-to-engine distance is too small incomparison with that for the selected aircraft 12A are stopped, i.e.,not accepted into the gate, the safety margin can always be kept at anacceptable level.

[0158] A situation in which the aircraft is at an angle relative to theDGS 10 will now be considered. As shown in FIG. 15A, a first aircraft12D can be aligned correctly relative to the DGS 10, whereas a secondaircraft 12D can deviate from the correct alignment by a yaw angle γ. Avery high-level description of the technique used in such a situation isthat the yaw angle of the aircraft is determined, and the profile isrotated to match that yaw angle.

[0159]FIG. 15 shows a flow chart of the technique. In step 1502, thepolar coordinates of the echoes returned from the aircraft are convertedto Cartesian coordinates. In step 1504, the yaw angle is calculated. Instep 1506, the echo profile is rotated. In step 1508, the IDcharacteristics are detected in the manner already described.

[0160] Step 1502 is carried out in the following manner. The echocoordinates received from the aircraft are converted from polarcoordinates (a_(j), r_(j)) to Cartesian coordinates (x_(j),y_(j)) withthe origin in the nose tip (α_(nose)r_(nose)) and with the y-axis alongthe line from the laser unit through the nose tip as follows:

[0161] x_(j)=r_(j) sin α_(j)

y _(j) =r _(j) cos α_(j) −r _(nose).

[0162] Step 1504 is carried out in a manner which will be explained withreference to FIGS. 16 and 17. FIG. 16 is a diagram showing the geometryof the regression lines on either side of the nose tip. FIG. 17 is aflow chart showing steps in the algorithm.

[0163] The algorithm is based on regression lines, calculated for echoesin a defined region behind the nose tip. If there are a sufficientnumber of echoes on both sides of the nose, then the yaw angle iscalculated from the difference in angle between the regression lines. Ifonly the regression line for one side of the nose can be calculated,e.g. due to the yaw angle, then the yaw angle is calculated from thedifference in angle between that regression line and the correspondingpart of the reference profile.

[0164] In step 1702, the echo coordinates are converted to Cartesiancoordinates (x_(j),y_(j)) in the manner described above. In step 1704,the approximate coordinates of the nose tip are calculated.

[0165] In step 1706, the echoes are screened in the following manner.Echoes not representative for the general shape of the echo picture areremoved before the angle of the echo picture is calculated. The echoscreening starts from the origin (the pointed out nose tip) and removesboth echoes if an echo at next higher angular step is at the same orshorter distance.

[0166] In step 1708, for each echo, the distance R_(ni) to the nose tipis calculated as

R _(nj) ={square root}{square root over (x_(j) ²+y_(j) ²)}.

[0167] In step 1710, for each side of the nose tip, the echoes areselected for which R_(nj) are larger than R_(min), which is a constant(in the order of 1-2 m) defined specifically for each aircraft type. Instep 1712, the following mean values are calculated:

x _(leftmean)=1/n _(left) ×Σx _(jleft) x _(rightman)=1/n _(right) ×Σx_(jright)

y _(leftmean)=1/n _(left) ×Σy _(jleft) y _(rightman)=1/n _(right) ×Σy_(jright)

x ² _(leftmean)=1/n _(left) ×Σx ² _(jleft) x ² _(rightman)=1/n _(right)×Σx ² _(jright)

xy _(leftmean)=1/n _(left)×Σ(x _(jleft) ×y _(jleft)) xy_(rightman)=1/n_(right)×Σ(x_(right) ×y _(right))

[0168] where n=the number of echoes >R_(min) on a respective side, andthe subscript right or left identifies the respective side to which aparticular quantity applies.

[0169] In step 1712, each regression line's angle v_(reg) to they axisis calculated as:$v_{reg} = {{arc}\quad \cot {\left\{ \frac{{xy}_{mean} - {x_{mean}y_{mean}}}{x_{mean}^{2} - \left( x_{mean} \right)^{2}} \right\}.}}$

[0170] The subscript mean should be read as leftmean or rightmean inaccordance with whether the angle is calculated on the left or rightside of the nose.

[0171] The yaw angle γ is calculated in the following manner. In step1714, it is determined whether the number n of echoes on both sides ofthe nose is greater than a predetermined value N, e.g., 5. If so, thenin step 1718, γ is calculated as

γ=(v _(regleft) +v _(regright))/2,

[0172] where v_(regleft) and v_(regright) are the angles calculated forthe left and right sides of the nose using the procedure of step 1712.On the other hand, if the n<N on one side of the nose, the referenceprofile is used for the calculation. In step 1720, the side and segmentof the profile are identified which correspond to the side where n>N. Instep 1722, the angle v_(refreg) is calculated for that segment using theprocedure of step 1712. Then γ is calculated in step 1718 asγ=(v_(refreg)−v_(reg)).

[0173] Once the yaw angle is calculated, then, in step 1506, the echoprofile is rotated accordingly. More specifically, the echo profile isconverted from one Cartesian coordinate system (x,y) to another (u, v)which has the same origin but is rotated by an angle equal to the yawangle γ, as shown in FIG. 18. The rotation of the echo profile will nowbe described with reference to FIGS. 18 and 19.

[0174] In step 1902, the approximate coordinates of the nose tip arecalculated. In step 1904, the echo coordinates are converted from polarto Cartesian coordinates (x_(i), y_(i)) with the nose tip as the originof the coordinate system. The technique for doing so has been describedabove. In step 1906, the echo coordinates are converted from the (x,y)coordinate system to the (u,v) coordinate system, as shown in FIG. 18,through the following formulae:

u _(i) =x _(i) cos γ+y_(i) sin γ;

v _(i) =−x _(i) sin γ+y_(i) cos γ.

[0175] The echo coordinates as thus rotated are used to identify theaircraft in the manner described above.

[0176] It will now be described how to set parameters defining centerlines (CL's), curved as well as straight, with reference to FIGS. 20-22.One docking system can handle several center lines with the technique tobe described.

[0177] The CL is specified as a piecewise linear curve, where α arecoordinates (α-sideways, l-lengthways) for the breaking points and areused as the defining parameters. The number of coordinates used ischosen with respect to required positioning accuracy. A straight CL isthus defined by the coordinates of two points (e.g. at the clip distanceand at a stop position). The number of coordinates required for a curvedCL depends on its radius.

[0178] The microprocessor 26 is used in the CL setting mode of step2002, in which the CL's are mapped in the microprocessor. A CL to bedefined is selected from a menu. One or more calibration poles withknown height and a top which is easily recognised in the calibrationpicture are placed on different positions on that CL. For each pole, theheight of the pole is typed in, and the top of the pole as appearing inthe calibration picture is clicked. The αand l coordinates for the poleare automatically entered in the table for that CL. The procedure isrepeated for each pole. The coordinates for the various poles areordered in the table by their l values. The number of poles neededdepends on the type of CL, with a straight CL needing only two and acurved CL needing more.

[0179] The calculation of the offset of the nose from the nose wheelwill now be discussed. The CL is normally given as the ideal nose-wheeltrack, but the guidance given to the aircraft is normally based on thenose position. That means, in case of a curved CL, that either the CLcoordinates must be converted to nose-coordinates, or the nose positionmust be converted to nose-wheel position. The latter is chosen, whichmeans that the yaw angle (v_(rot)) of the aircraft is determined in step2004 in the manner described above.

[0180] The nose-wheel position (α_(w),l_(w)) is calculated in step 2006as follows:

α_(w)>>α_(n) +l _(nw)×sin v_(rot)/(l _(n) +l _(nw×cos) v _(rot))  (inrad.)

l _(w) >>l _(n) +l _(nw)×cos v_(rot)

[0181] where

[0182] α_(jp) _(l) _(n): measured position of the nose;

[0183] l_(nw): nose-wheel distance; and

[0184] v_(rot): estimated yaw angle of aircraft.

[0185] The offset of the nose wheel from the CL is calculated in step2008 as follows:

Offet=α_(i)−α_(w)+(l _(w) +l _(i))(α_(i+l)−α_(i))/(l _(i+l) −l _(i))

[0186] where

[0187] α_(jp) l_(i) is the CL-coordinate pair with l_(i)-value justbelow l_(w); and

[0188] α_(j+l), l_(i+l) is the CL-coordinate pair with l_(i)-value justabove l_(w).

[0189] The calculations of step 2006 will now be explained withreference to FIG. 21, in which:

[0190] l_(nw): nose-wheel distance

[0191] v: estimated yaw angle of aircraft

[0192] x: estimated sideways position of nose-wheel

α_(w)>>α_(n) +x/(l_(n) +l _(nw)×cos v)  (in rad.)

l_(w) >>l _(n) +l _(nw)×cos v

x=l_(nw)×sin v

[0193] The calculations of step 2008 will now be explained withreference to FIG. 22, in which x₀/y₀ represents the estimated positionof the nose wheel and x_(i)/y_(i) represents the breaking points in thepiecewise-linear model of a curved CL. The “real” offset from the CL isthe distance measured in a right angle to the CL. An approximation ofthat distance is the distance measured in a right angle to the laserbeam from the docking system. That distance corresponds to the value(x_(m)−x₀) in FIG. 22. As the absolute value of the offset not isimportant, that approximation is used. From FIG. 22, it follows that

Offset=(x _(m) −x ₀)=x_(i) −x ₀+(y ₀ −y _(i))(x _(i+1) −x _(i))/(y_(i+1) −y _(i)).

[0194] While a preferred embodiment of the present invention has beenset forth in detail above, those skilled in the art will readilyappreciate that other embodiments can be realized within the scope ofthe invention. For example, while the aircraft criterion phase 408 isdisclosed as using the ratio Vi/(Vi+Vo), the difference Vi−Vo could beused instead. Also, the specific numerical ranges disclosed above shouldbe considered to be illustrative rather than limiting. Those skilled inthe art will be able to derive other numerical ranges as needed to adaptthe invention to other models of aircraft or to the specific needs ofvarious airports. Furthermore, while regression lines are a usefultechnique for determining the yaw angle, any other technique can beused. Therefore, the present invention should be construed as limitedonly by the appended claims. TABLE I 42

78.25

77.5 . . . 23 44

5 5 5.6 . . . 10 45

1 2 3 . . . 50 0 xx xx xx . . . xx 1 xx xx xx . . . xx 2 xx xx xx . . .xx 3 xx xx xx . . . xx 4 xx xx xx . . . xx 5 xx xx xx . . . xx 6 xx xxxx . . . xx 7 xx xx xx . . . xx 8 xx xx xx . . . xx 9 xx xx xx . . . xx. . . 50 xx xx xx . . . xx

[0195] TABLE II 91 92 93 94 95 96 97

i l_(i) l_(ki) d_(i) a_(i) a_(e) Note 1 xx xx xx xx xx xx 2 xx xx xx xxxx xx 3 xx xx xx xx xx xx 4 xx xx xx xx xx xx 5 xx xx xx xx xx xx 6 xxxx xx xx xx xx . . . . 50 xx xx xx xx xx xx . . . 100 xx xx xx xx xx xx

What is claimed is:
 1. A system for determining whether a detectedobject is a known object, the known object having a known profile andalso having a known feature at a known location, the system comprising:projecting means for projecting light pulses onto the detected object;collecting means for collecting light pulses reflected off the detectedobject and for detecting a shape of the detected object in accordancewith the light pulses; comparing means for comparing the detected shapewith a profile corresponding to the known shape and for determiningwhether the detected shape corresponds to the known shape; andidentifying means for identifying whether the detected object is theknown object by determining whether the detected object has the knownfeature at the known location.
 2. The system of claim 1, wherein: forthe known object, an inner volume is defined so as to contain the knownfeature, and an outer volume is defined so as not to contain the knownfeature; the identifying means determines whether the detected objecthas the known feature in the known location in accordance with a numberof light pulses reflected from within the inner volume and a number oflight pulses reflected from within the outer volume.
 3. The system ofclaim 2, wherein the outer volume is defined to surround the innervolume.
 4. The system of claim 2, wherein the identifying meansdetermines whether the detected object has the known feature in theknown location in accordance with whether Vi/(Vi+Vo)>T, where: Vi=thenumber of light pulses reflected from the inner volume; Vo=the number oflight pulses reflected from the outer volume; and T=a predeterminedthreshold value.
 5. The system of claim 4, wherein T=0.7.
 6. The systemof claim 2, wherein the identifying means controls the projecting meansto project light pulses into the inner volume and the outer volume. 7.The system of claim 1, wherein: the known object comprises a nose with aknown nose height; and the identifying means further identifies whetherthe detected object is the known object by detecting a nose height ofthe detected object and comparing the detected nose height to the knownnose height.
 8. The system of claim 7, wherein the identifying meanscompares the detected nose height to the known nose height by taking adifference between the detected nose height and the known nose height.9. The system of claim 8, wherein the identifying means identifies thedetected object as the known object only if the difference is less thanor equal to a threshold difference.
 10. The system of claim 9, whereinthe threshold difference is 0.5 m.
 11. The system of claim 1, whereinthe comparing means determines a yaw angle of the detected object. 12.The system of claim 11, wherein the comparing means rotates the profilecorresponding to the known shape by an angle equal to the yaw angle. 13.A method for determining whether a detected object is a known object,the known object having a known profile and also having a known featureat a known location, the method comprising: (a) projecting light pulsesonto the detected object; (b) collecting light pulses reflected off thedetected object and for detecting a shape of the detected object inaccordance with the light pulses; (c) comparing the detected shape witha profile corresponding to the known shape and for determining whetherthe detected shape corresponds to the known shape; and (d) identifyingwhether the detected object is the known object by determining whetherthe detected object has the known feature at the known location.
 14. Themethod of claim 13, wherein: for the known object, an inner volume isdefined so as to contain the known feature, and an outer volume isdefined so as not to contain the known feature; said step of identifyingcomprises determining whether the detected object has the known featurein the known location in accordance with a number of light pulsesreflected from within the inner volume and a number of light pulsesreflected from within the outer volume.
 15. The method of claim 14,wherein the outer volume is defined to surround the inner volume. 16.The method of claim 14, wherein said step of identifying comprisesdetermining whether the detected object has the known feature in theknown location in accordance with whether Vi/(Vi+Vo)>T, where: Vi=thenumber of light pulses reflected from the inner volume; Vo=the number oflight pulses reflected from the outer volume; and T=a predeterminedthreshold value.
 17. The method of claim 16, wherein T=0.7.
 18. Themethod of claim 14, wherein said step of identifying comprisescontrolling said step of projecting to project light pulses into theinner volume and the outer volume.
 19. The method of claim 13, wherein:the known object comprises a nose with a known nose height; and saidstep of identifying comprises further identifying whether the detectedobject is the known object by detecting a nose height of the detectedobject and comparing the detected nose height to the known nose height.20. The method of claim 19, wherein said step of identifying comprisescomparing the detected nose height to the known nose height by taking adifference between the detected nose height and the known nose height.21. The method of claim 20, wherein said step of identifying identifiesthe detected object as the known object only if the difference is lessthan or equal to a threshold difference.
 22. The method of claim 21,wherein the threshold difference is 0.5 m.
 23. The method of claim 13,wherein said step of comparing comprises determining a yaw angle of thedetected object.
 24. The method of claim 23, wherein said step ofcomparing further comprises rotating the profile corresponding to theknown shape by an angle equal to the yaw angle.
 25. A system fordetermining a yaw angle of a detected object, the system comprising:projecting means for projecting light pulses onto the detected object;collecting means for collecting light pulses reflected off the detectedobject and for detecting a shape of the detected object in accordancewith the light pulses; and angle determining means for determining theyaw angle from the shape detected by the collecting means.
 26. Thesystem of claim 25, wherein the detected object comprises a nose havinga nose tip, and wherein the angle determining means determines the yawangle from a portion of the shape which is adjacent to the nose tip. 27.The system of claim 26, wherein the nose has a left side and a rightside relative to the nose tip, and wherein the angle determining meansdetermines a regression line on at least one of the left side and theright side and determines the yaw angle in accordance with theregression line.
 28. A method of determining a yaw angle of a detectedobject, the method comprising: projecting light pulses onto the detectedobject; collecting light pulses reflected off the detected object anddetecting a shape of the detected object in accordance with the lightpulses; and determining the yaw angle from the shape detected by thecollecting means.
 29. The method of claim 28, wherein the detectedobject comprises a nose having a nose tip, and wherein said step ofdetermining comprises determining the yaw angle from a portion of theshape which is adjacent to the nose tip.
 30. The method of claim 29,wherein the nose has a left side and a right side relative to the nosetip, and wherein said step of determining comprises determining aregression line on at least one of the left side and the right side anddetermining the yaw angle in accordance with the regression line.
 31. Asystem for determining whether a vehicle is following a center line, thevehicle having a nose and a wheel, the system comprising: a storagedevice for storing (i) coordinates representing a path of the centerline and (ii) a distance between the nose and the wheel; a detectingdevice for detecting (i) a position of the nose and (ii) a yaw angle ofthe vehicle; and a calculating device for calculating (i) a position ofthe wheel, from the position of the nose, the yaw angle detected by thedetecting device and the distance stored in the storage device, and (ii)an offset of the wheel from the center line, from the coordinates storedin the storage device and the position of the wheel.
 32. A method ofdetermining whether a vehicle is following a center line, the vehiclehaving a nose and a wheel, the method comprising: storing coordinatesrepresenting a path of the center line; storing a distance between thenose and the wheel; detecting a position of the nose; detecting a yawangle of the vehicle; calculating a position of the wheel, from theposition of the nose, the yaw angle detected by the detecting device andthe distance stored in the storage device; and calculating an offset ofthe wheel from the center line, from the coordinates stored in thestorage device and the position of the wheel.